EP0578951B1 - Jet engine - Google Patents

Jet engine Download PDF

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Publication number
EP0578951B1
EP0578951B1 EP93108556A EP93108556A EP0578951B1 EP 0578951 B1 EP0578951 B1 EP 0578951B1 EP 93108556 A EP93108556 A EP 93108556A EP 93108556 A EP93108556 A EP 93108556A EP 0578951 B1 EP0578951 B1 EP 0578951B1
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EP
European Patent Office
Prior art keywords
nozzle
extension piece
thrust
engine
wall
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
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EP93108556A
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German (de)
French (fr)
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EP0578951A1 (en
Inventor
Joachim Dipl.-Ing. Kretschmer
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Airbus Defence and Space GmbH
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Daimler Benz Aerospace AG
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Publication of EP0578951A1 publication Critical patent/EP0578951A1/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/28—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow
    • F02K1/30—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow for varying effective area of jet pipe or nozzle
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/06—Varying effective area of jet pipe or nozzle
    • F02K1/09—Varying effective area of jet pipe or nozzle by axially moving an external member, e.g. a shroud
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
    • F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
    • F02K7/16—Composite ram-jet/turbo-jet engines

Definitions

  • the invention relates to a jet engine, in particular a hypersonic combination engine, according to the preamble of claims 1 and 8, respectively.
  • hypersonic combination engines of the type claimed which work at low flight speeds up to about 3 Mach as a turbine jet engine and in the upper speed range between 3 and about 7 Mach as a ramjet engine and therefore Need thrusters that have a very wide range of variation both in terms of the narrowest cross-sectional area and in terms of the expansion ratio and for this purpose in addition to a mushroom-shaped central body that can be adjusted axially between the inlet and outlet cross-sections of the divergent outer wall of the nozzle at the nozzle outlet end to change the nozzle neck cross-section have swiveling tailgates, of which the upper one, in the ram-jet mode in the fully swiveled-out position, bears against an expansion ramp on the missile assigned to the engine and, in turbine jet mode, when the tailgates are swiveled into a position narrowing the hot gas flow channel, an air channel is freed via which the boundary layer air is blown from the engine inlet into the
  • the object of the invention is to design a jet engine of the type mentioned in such a way that the geometry of the thruster-limited hot gas flow channel adapts in a structurally simple manner with high propulsion efficiency to greatly differing operating conditions, such as occur in particular in the hypersonic combination drives described leaves.
  • the axially movable, bell-shaped, divergent thrust nozzle extension makes a highly variable cross-sectional profile of the hot gas flow channel delimited by the nozzle contour both in terms of the outlet area and in terms of the area of the nozzle neck received, so that the thrust nozzle easily with high Propulsion efficiency can be set to greatly differing nozzle pressure ratios and hot gas flows, for example when switching from turbo to ramjet drive, and on the other hand in connection with the annular secondary air injection at the outlet end of the fixed nozzle outer wall, which is retracted by the hot gas jet limitation of the thrust nozzle extension the dead water effect, which is particularly pronounced in the trans sound area, is effectively reduced, thereby significantly reducing the rear drag and the propulsion-reducing vacuum zone on the underside of the engine, with the further structural simplification that the axially movable nozzle extension part not only extends the nozzle outer wall in the extended
  • the nozzle extension part in the retracted state for regulating the air duct outlet cross section can be adjusted axially to a limited extent in order to achieve the greatest possible additional thrust gain through the air injection.
  • an engine-fixed, circumferentially closed distributor ring encompassing the outer wall of the nozzle, with an annular gap which is open in the direction of the nozzle outlet end and is sealed by this in the retracted state of the adjustable nozzle extension part.
  • seal between the distributor ring and the nozzle extension part with regard to the preferred controllability of the air duct cross-section is expediently carried out according to claim 4 by an effective sliding seal in the retracted state of the extension part within a limited axial stroke and the distributor ring in an installation-friendly manner according to claim 5 preferably in the region of the smallest nozzle wall outer diameter is arranged.
  • shut-off devices are expediently provided according to claim 6 upstream of the air outlet channel or distributor ring.
  • the boundary layer air is used in a particularly preferred manner for secondary air injection at the divergent end of the nozzle outer wall.
  • the central body with its largest cross section is arranged in a particularly preferred manner according to claim 9 in the extended position of the divergent nozzle extension between the narrowest cross section of the engine-fixed nozzle outer wall and the exit plane of the nozzle extension.
  • a further, controlled expansion of the thrust jet is achieved on the downstream side of the extended thrust nozzle extension, preferably by means of an expansion ramp fixed to the missile on one side and delimiting the thrust jet.
  • a hypersonic aircraft which is attached to the underside of a hypersonic aircraft 2, not shown in detail, consists essentially of an air inlet 4, an internal turbo engine 6, a ramjet engine consisting of a ram air duct 8 and a combustion chamber 10 with injectors 12 14 and a thruster 16.
  • the turbo engine 6 comprises a turbo inlet duct 22 which can be closed by a switchover element 20 and which has a low-pressure and a high-pressure compressor 18, 24, a combustion chamber 26, a turbine 28 and a turbo outlet duct 32 which can be closed by a ring slide 30.
  • turbo mode which is switched on from the start up to a switchover number of approximately 3 with nozzle pressure ratios between approximately 4 and 40, the air flowing in via the inlet 4 reaches the turbo inlet duct 22 via the then open switchover element 20, where it is compressed by means of the compressors 18, 24 compressed and then burned in the combustion chamber 26 together with liquid hydrogen.
  • the hot gas flow is introduced into the ram air duct 8 via the ring slide 30, which is also open, and after passing through the injectors 12 and the combustion chamber 10, which is operated in turbo mode from about 0.9 as an afterburner with hydrogen as fuel , relaxed in the nozzle 16 scraping.
  • a closure flap 34 is shown in FIG.
  • the switching member 20 and the ring valve 30 as well as the closure flap 34 are closed, so that the turbo engine 6 and the boundary layer channel 36 is deactivated, and the entire air flow is impact-compressed by means of an adjustable ramp arrangement 38 and flows via the ram air channel 8 into the combustion chamber 10, where it is burned together with hydrogen supplied via the injectors 12 and then expands in the hot gas flow channel delimited by the thrust nozzle 16 to produce thrust.
  • the design of the thrust nozzle 16 and the different cross-sectional configurations of the hot gas flow channel delimited by it in the different flight conditions are explained in more detail with reference to FIGS. 2 to 4.
  • the main components of the thrust nozzle 16 include an engine-fixed, convergent-divergent, rotationally symmetrical nozzle outer wall 42 connected to the cylindrical combustion chamber wall 40, a mushroom-shaped central body 44, which is axially adjustable on a central, engine-fixed carrier 46, an axially movable, divergent, also rotationally symmetrical Thrust nozzle extension part 48 and a cold air distributor ring 50, which surrounds the nozzle outer wall 42 in the region of the convergent wall section, with an air inlet 52 connected to the rear end of the boundary layer channel 36.
  • the central body 44 with its largest cross section D lies between the narrowest point and the outlet end 54 of the outer wall 42 of the nozzle and, together with it, delimits a convergent-divergent hot gas flow channel 56 an annular nozzle neck cross section h 1 , while the nozzle extension part 48, which is not required in this flight state for expanding the hot gas flow channel 56, is retracted into the retracted position and in this the outer ring wall of an air outlet channel 58 adjoining the distributor ring 50, which forms the surrounds the divergent end of the nozzle outer wall 42 in an annular manner, the extension part 48 interacting in this position via a sliding seal 60 with an annular gap 62 of the distributor ring 50 which is open in the direction of the nozzle outlet end.
  • the boundary layer air separated in this state on the inlet side is blown in via the air outlet duct 58 delimited between the outer wall 42 of the nozzle and the extension part 48 over the entire outer circumference of the outlet end 54, which not only achieves an additional thrust effect, but above all also the dead water zone 64, which is otherwise in this flight state between the - shown in dashed lines in the figures - Shear jet downstream of the nozzle outer wall 42 and the ambient air flow (shown in phantom) forms, effectively reduced.
  • FIG. 3 shows the thrust nozzle 16 in turbo mode close to the Mach Mach number.
  • the central body 44 is adjusted on the carrier 46 with the largest central body diameter D up to the outlet cross section of the fixed nozzle outer wall 42, so that the nozzle neck cross section is increased to h 2 .
  • the nozzle extension part 48 is still in the retracted state, but is limited axially adjustable due to the sliding sealing effect without loss of sealing with respect to the annular gap 62 and thus enables regulation of the effective flow cross section of the air outlet duct 58, via which according to FIG. 3 continues - but now essentially reduced flow cross-section - boundary layer air is blown out into the dead water zone 64.
  • Fig. 4 shows the thrust nozzle 16 in the upper half shortly after switching to the ram air drive.
  • the nozzle extension part 48 is shifted from the retracted state interacting with the sliding seal 60 to the extended position, in which it adjoins the outlet end 54 of the nozzle outer wall 42 in the same shape, whereby the expansion ratio of the nozzle outer contour clearly to one of the outlet surface A of the extension part 48 in relation to smallest cross-sectional area E of the nozzle outer wall 42 corresponding value is increased.
  • the thrust jet extends to the ambient flow, so that the disturbing effect of a dead water zone at the outlet end of the hot gas flow channel is eliminated and the air outlet channel 58 is no longer required for secondary air injection.
  • the boundary layer separation is dispensed with, and therefore the closure flap 34 is pivoted into the closed position shown in FIG. 1.
  • the central body 44 is in the rear end position, in which its largest cross-sectional area D lies in the exit plane A of the divergent extension part 48 and the nozzle neck cross section h 3 of the hot gas flow channel 56 therefore reaches its maximum value.
  • the central body 44 is moved forward with a continuous reduction in the nozzle neck cross section (and the opposite increase in the hot gas outlet area at the divergent end of the nozzle extension part 48) until the largest central body cross section D in the plane E is smallest Is the outer diameter of the nozzle wall and therefore the hot gas flow channel 56 delimited by the thrust nozzle reaches its minimum nozzle neck cross section h 4 and the maximum expansion ratio (lower half of FIGS. 4; and 1).
  • an expansion ramp 66 of the same shape is arranged at the tail end of the missile 2 for one-sided thrust jet limitation.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Jet Pumps And Other Pumps (AREA)
  • Control Of Turbines (AREA)

Description

Die Erfindung bezieht sich auf ein Strahltriebwerk, insbesondere ein Hyperschall-Kombinationstriebwerk, nach dem Oberbegriff des Patentanspruchs 1 bzw. 8.8. The invention relates to a jet engine, in particular a hypersonic combination engine, according to the preamble of claims 1 and 8, respectively.

Aus der US-PS 4 527 388 ist ein Mantelstrom-Triebwerk mit einer konvergenten Schubdüse und einem zur Veränderung des Düsenhalsquerschnitts im Düseninneren axial verstellbaren, pilzförmigen Zentralkörper bekannt, bei dem der um das Grundtriebwerk herumgeführte äussere Kaltluftstrom dem die Schubdüse durchsetzenden Heißgasstrom über einen am Düsenaustrittsende angeordneten Luftauslaßkanal mit einer die Schubdüse konvergent umschließenden äußeren Begrenzungswand zugemischt wird, welche zur Regulierung des wirksamen Luftkanalquerschnitts und damit des Kaltluft-/Heißgasstromverhältnisses unabhängig vom Zentralkörper axial verstellbar ist. Ein derartiges Mantelstromtriebwerk mit einer am Austrittsende konvergenten, zur Anpassung an unterschiedliche Zweistrom-Mengenverhältnisse verstellbaren Schubdüse für einen hinsichtlich Flughöhe und -geschwindigkeit begrenzten Einsatzbereich ist nicht Gegenstand der Erfindung.From US Pat. No. 4,527,388, a turbofan jet engine with a convergent thrust nozzle and a mushroom-shaped central body which is axially adjustable for changing the nozzle neck cross section in the nozzle interior is known, in which the outer cold air stream which is guided around the basic engine and the hot gas stream passing through the thrust nozzle via an end of the nozzle outlet arranged air outlet channel is mixed with an outer boundary wall convergingly enclosing the thrust nozzle, which is axially adjustable independently of the central body to regulate the effective air channel cross section and thus the cold air / hot gas flow ratio. Such a turbofan engine with a convergent at the outlet end, adjustable to adapt to different two-flow quantity ratios, for an application area limited in terms of altitude and speed is not the subject of the invention.

Ferner sind aus der DE 40 10 471 A1 und der DE 40 12 212 A1 Hyperschall-Kombinationstriebwerke der beanspruchten Art bekannt, die bei niedrigen Fluggeschwindigkeiten bis zu etwa 3 Mach als Turbinenstrahltriebwerk und im oberen Geschwindigkeitsbereich zwischen 3 und etwa 7 Mach als Staustrahltriebwerk arbeiten und daher Schubdüsen benötigen, die sowohl hinsichtlich der engsten Querschnittsfläche als auch hinsichtlich des Erweiterungsverhältnisses über eine sehr hohe Variationsbreite verfügen und zu diesem Zweck zusätzlich zu einem pilzförmigen Zentralkörper, der zur Veränderung des Düsenhalsquerschnitts axial zwischen dem Ein- und Austrittsquerschnitt der divergenten Düsenaußenwand verstellbar ist, am Düsenaustrittsende schwenkbare Heckklappen besitzen, von denen die obere sich im Staustrahlbetrieb in der voll ausgeschwenkten Lage an eine Expansionsrampe an dem dem Triebwerk zugeordneten Flugkörper anlegt und im Turbinenstrahlbetrieb, wenn die Heckklappen auf eine den Heißgas-Strömungskanal verengende Position eingeschwenkt sind, einen Luftkanal freisteuert, über den die Grenzschichtluft vom Triebwerkseinlauf in den Heißgasstrahl im Bereich der Expansionsrampe stromabwärts der oberen Heckklappe eingeblasen wird. Problematisch bei derartigen Triebwerken ist zum einen der hohe mechanische Bau- und vor allem Dichtaufwand für die verschwenkbaren Schubdüsenklappen und zum anderen die gerade in kritischen Flugzuständen ausgeprägte Vortriebseinbuße. So ergibt sich im Transschallbetrieb eine höchst störende Unterdruckzone an der Triebwerksunterseite und Zunahme des Heckwiderstands, die ihre Ursache in einer sich im eingeschwenkten Zustand der Düsenklappen am divergenten Ende der Schubdüsenaußenwand aufgrund einer Ablösung der Umgebungsströmung ausbildenden Totwasserregion haben, und ferner kommt es in der Start- und/oder der Umschaltphase vom Turbo- auf den Staustrahlbetrieb zu deutlichen Schubverlusten, weil die erzielbaren Querschnittsänderungen des Heißgas-Strömungskanals nicht für eine schuboptimale Anpassung an den sich drastisch ändernden Heißgas-Volumenstrom ausreichen.Furthermore, from DE 40 10 471 A1 and DE 40 12 212 A1 hypersonic combination engines of the type claimed are known, which work at low flight speeds up to about 3 Mach as a turbine jet engine and in the upper speed range between 3 and about 7 Mach as a ramjet engine and therefore Need thrusters that have a very wide range of variation both in terms of the narrowest cross-sectional area and in terms of the expansion ratio and for this purpose in addition to a mushroom-shaped central body that can be adjusted axially between the inlet and outlet cross-sections of the divergent outer wall of the nozzle at the nozzle outlet end to change the nozzle neck cross-section have swiveling tailgates, of which the upper one, in the ram-jet mode in the fully swiveled-out position, bears against an expansion ramp on the missile assigned to the engine and, in turbine jet mode, when the tailgates are swiveled into a position narrowing the hot gas flow channel, an air channel is freed via which the boundary layer air is blown from the engine inlet into the hot gas jet in the area of the expansion ramp downstream of the upper tailgate. Problems with such engines are, on the one hand, the high mechanical construction and, above all, the sealing effort for the pivotable throttle valve flaps, and on the other hand the loss of propulsion, which is particularly pronounced in critical flight conditions. This results in a highly disturbing negative pressure zone on the underside of the engine and an increase in the rear drag in trans-sound mode, which is caused by a dead water region which forms when the nozzle flaps are swiveled in at the divergent end of the thrust nozzle outer wall due to a detachment of the surrounding flow, and further occurs in the start and / or the switchover phase from turbo to ramjet operation to significant thrust losses because the achievable cross-sectional changes of the hot gas flow channel are not sufficient for an optimal thrust adjustment to the drastically changing hot gas volume flow.

Aufgabe der Erfindung ist es, ein Strahltriebwerk der eingangs genannten Art so auszubilden, daß sich die Geometrie des schubdüsenbegrenzten Heißgas-Strömungskanals auf konstruktiv einfache Weise mit hohem Vortriebs-Wirkungsgrad an stark differierende Betriebsverhältnisse, wie sie insbesondere bei den geschilderten Hyperschall-Kominationsantrieeen auftreten, anpassen läßt.The object of the invention is to design a jet engine of the type mentioned in such a way that the geometry of the thruster-limited hot gas flow channel adapts in a structurally simple manner with high propulsion efficiency to greatly differing operating conditions, such as occur in particular in the hypersonic combination drives described leaves.

Diese Aufgabe wird erfindungsgemäß durch das im Patentanspruch 1 und/oder 8 gekennzeichnete Strahltriebwerk gelöst.This object is achieved by the jet engine characterized in claim 1 and / or 8.

Erfindungsgemäß wird durch die axial verfahrbare, glockenförmige, divergente Schubdüsenverlängerung einerseits in Verbindung mit der beanspruchten Zentralkörper-Verstellbarkeit auf baulich einfache, leckagearme Weise ein sowohl hinsichtlich der Austritts- als auch hinsichtlich der Düsenhals-Flächengröße hochgradig variabler Querschnittsverlauf des von der Düsenkontur begrenzten Heißgas-Strömungskanals erhalten, so daß sich die Schubdüse problemlos mit hohem Vortriebswirkungsgrad auf stark unterschiedliche Düsendruckverhältnisse und Heißgasströme, etwa bei der oben erwähnten Umschaltung vom Turbo- auf den Staustrahlantrieb, einstellen läßt, und andererseits in Verbindung mit der ringförmigen, in der von der Heißgas-Strahlbegrenzung zurückgezogenen Lage der Schubdüsenverlängerung zugeschalteten Sekundärlufteinblasung am Austrittsende der festen Düsenaußenwand der dort insbesondere im Transschallbereich ausgeprägte Totwasser-Effekt wirksam abgebaut und dadurch der Heckwiderstand und die vortriebsmindernde Unterdruckzone auf der Triebwerksunterseite deutlich reduziert, mit der weiteren baulichen Vereinfachung, daß das axial verfahrbare Düsenverlängerungsteil nicht nur in der Ausfahrposition die Erweiterung der Düsenaußenwand, sondern zusätzlich im eingefahrenen Zustand auch die äußere Umgrenzung des Luftkanals und Steuerung der Sekundärlufteinblasung übernimmt.According to the invention, the axially movable, bell-shaped, divergent thrust nozzle extension, on the one hand, in conjunction with the claimed adjustability of the central body in a structurally simple, low-leakage manner, makes a highly variable cross-sectional profile of the hot gas flow channel delimited by the nozzle contour both in terms of the outlet area and in terms of the area of the nozzle neck received, so that the thrust nozzle easily with high Propulsion efficiency can be set to greatly differing nozzle pressure ratios and hot gas flows, for example when switching from turbo to ramjet drive, and on the other hand in connection with the annular secondary air injection at the outlet end of the fixed nozzle outer wall, which is retracted by the hot gas jet limitation of the thrust nozzle extension the dead water effect, which is particularly pronounced in the trans sound area, is effectively reduced, thereby significantly reducing the rear drag and the propulsion-reducing vacuum zone on the underside of the engine, with the further structural simplification that the axially movable nozzle extension part not only extends the nozzle outer wall in the extended position, but also in the retracted position Condition also takes over the outer boundary of the air duct and control of the secondary air injection.

Im Hinblick auf eine weitere Verbesserung der Vortriebswirkung ist in vorteilhafter Ausgestaltung der Erfindung gemäß Anspruch 2 das Düsenverlängerungsteil im eingefahrenen Zustand zur Regelung des Luftkanal-Auslaßquerschnitts begrenzt axial verstellbar, um durch die Lufteinblasung einen möglichst großen zusätzlichen Schubgewinn zu erzielen. Aus Gründen einer gleichmäßigen Sekundärluftverteilung und baugünstigen Abdichtung ist ferner gemäß Anspruch 3 zur Sekundärluftzufuhr zum Luftauslaßkanal vorzugsweise ein die Düsenaußenwand umgreifender, triebwerksfester, in Umfangsrichtung geschlossener Verteilerring mit einem in Richtung des Düsenaustrittsendes offenen, im eingefahrenen Zustand des verstellbaren Düsenverlängerungsteils von diesem dichtend umschlossenen Ringspalt vorgesehen, wobei die Abdichtung zwischen dem Verteilerring und dem Düsenverlängerungsteil im Hinblick auf die bevorzugte Regulierbarkeit des Luftkanalquerschnitts zweckmäßigerweise gemäß Anspruch 4 durch eine im eingefahrenen Zustand des Verlängerungsteils innerhalb eines begrenzten Axialhubs wirksame Gleitdichtung erfolgt und der Verteilerring in einbaugünstiger Weise gemäß Anspruch 5 bevorzugt im Bereich des geringsten Düsenwandaußendurchmessers angeordnet ist.With a view to a further improvement of the propulsion effect, in an advantageous embodiment of the invention, the nozzle extension part in the retracted state for regulating the air duct outlet cross section can be adjusted axially to a limited extent in order to achieve the greatest possible additional thrust gain through the air injection. For the sake of a uniform secondary air distribution and a structurally advantageous seal, it is further preferred, according to claim 3, for secondary air supply to the air outlet duct, an engine-fixed, circumferentially closed distributor ring encompassing the outer wall of the nozzle, with an annular gap which is open in the direction of the nozzle outlet end and is sealed by this in the retracted state of the adjustable nozzle extension part. wherein the seal between the distributor ring and the nozzle extension part with regard to the preferred controllability of the air duct cross-section is expediently carried out according to claim 4 by an effective sliding seal in the retracted state of the extension part within a limited axial stroke and the distributor ring in an installation-friendly manner according to claim 5 preferably in the region of the smallest nozzle wall outer diameter is arranged.

Zum Abschalten der in der Ausfahrposition des Düsenverlängerungsteils nicht benötigten Sekundärlufteinblasung sind gemäß Anspruch 6 zweckmäßigerweise stromaufwärts des Luftauslaßkanals bzw. Verteilerrings Absperrorgane vorgesehen.To shut off the secondary air injection which is not required in the extended position of the nozzle extension part, shut-off devices are expediently provided according to claim 6 upstream of the air outlet channel or distributor ring.

Bei Triebwerken, die mit einer Grenzschichtabsaugung am Triebwerkseinlauf in Flugzuständen ohne Heißgaskanal-Verlängerung mittels des ausfahrbaren Schubdüsenteils arbeiten, insbesondere Hyperschalltriebwerken im Turbobetrieb, wird gemäß Anspruch 7 in besonders bevorzugter Weise die Grenzschichtluft zur Sekundärlufteinblasung am divergenten Ende der Düsenaußenwand verwendet.In engines that work with a boundary layer suction at the engine inlet in flight conditions without extending the hot gas duct by means of the extendable thrust nozzle part, in particular hypersonic engines in turbo operation, the boundary layer air is used in a particularly preferred manner for secondary air injection at the divergent end of the nozzle outer wall.

Um die Variationsbreite der erzielbaren Düsenhalsquerschnittsänderungen weiter zu erhöhen, ist der Zentralkörper mit seinem größten Querschnitt in besonders bevorzugter Weise gemäß Anspruch 9 in der Ausfahrposition der divergenten Düsenverlängerung zwischen dem engsten Querschnitt der triebwerksfesten Düsenaußenwand und der Austrittsebene der Düsenverlängerung verstellbar angeordnet.In order to further increase the range of variation of the achievable nozzle neck cross section changes, the central body with its largest cross section is arranged in a particularly preferred manner according to claim 9 in the extended position of the divergent nozzle extension between the narrowest cross section of the engine-fixed nozzle outer wall and the exit plane of the nozzle extension.

Schließlich wird gemäß Anspruch 10 eine weitere, gesteuerte Expansion des Schubstrahls abströmseitig der ausgefahrenen Schubdüsenverlängerung vorzugsweise durch eine den Schubstrahl einseitig begrenzende, flugkörperfeste Expansionsrampe erzielt.Finally, according to claim 10, a further, controlled expansion of the thrust jet is achieved on the downstream side of the extended thrust nozzle extension, preferably by means of an expansion ramp fixed to the missile on one side and delimiting the thrust jet.

Die Erfindung wird nunmehr anhand eines Ausführungsbeispiels in Verbindung mit den Zeichnungen näher erläutert. Es zeigen in stark schematisierter Darstellung:

Fig. 1
ein Hyperschall-Kombinationstriebwerk im Längsschnitt mit ausgefahrener Schubdüsenverlängerung und der Zentralkörper-Position kleinsten Düsenhalsquerschnitts im Hyperschallbetrieb;
Fig. 2
eine vergrößerte Darstellung der Schubdüse im Unterschall-Turbobetrieb;
Fig. 3
eine der Fig. 2 entsprechende Darstellung der oberen Schubdüsenhälfte im Überschallturbobetrieb; und
Fig. 4
die Schubdüse im Stauluftbetrieb, in der oberen Figurenhälfte nahe der Umschalt-Machzahl und in der unteren Figurenhälfte im Hyperschallflug.
The invention will now be explained in more detail using an exemplary embodiment in conjunction with the drawings. They show in a highly schematic representation:
Fig. 1
a hypersonic combination engine in longitudinal section with extended thrust nozzle extension and the central body position of the smallest nozzle neck cross section in hypersonic operation;
Fig. 2
an enlarged view of the thruster in subsonic turbo mode;
Fig. 3
one of FIG 2 corresponding representation of the upper thrust half in supersonic turbo mode. and
Fig. 4
the thruster in ram air mode, in the upper half of the figure near the switch Mach number and in the lower half of the figure in hypersonic flight.

Das in Fig. 1 gezeigte Hpperschall-Kombinationstriebwerk, welches auf der Unterseite eines nicht näher dargestellten Hyperschallflugzeugs 2 befestigt ist, besteht im wesentlichen aus einem Lufteinlaß 4, einem innenliegenden Turbotriebwerk 6, einem aus einem Stauluftkanal 8 und einer Brennkammer 10 mit Einspritzvorrichtungen 12 bestehenden Staustrahltriebwerk 14 und einer Schubdüse 16.1, which is attached to the underside of a hypersonic aircraft 2, not shown in detail, consists essentially of an air inlet 4, an internal turbo engine 6, a ramjet engine consisting of a ram air duct 8 and a combustion chamber 10 with injectors 12 14 and a thruster 16.

Das Turbotriebwerk 6 umfaßt einen durch ein Umschaltorgan 20 verschließbaren Turboeinlaßkanal 22 mit einem Niederdruck- und einem Hochdruckverdichter 18, 24, eine Brennkammer 26, eine Turbine 28 und einen durch einen Ringschieber 30 verschließbaren Turboauslaßkanal 32.The turbo engine 6 comprises a turbo inlet duct 22 which can be closed by a switchover element 20 and which has a low-pressure and a high-pressure compressor 18, 24, a combustion chamber 26, a turbine 28 and a turbo outlet duct 32 which can be closed by a ring slide 30.

Im Turbobetrieb, der vom Start bis zu einer Umschaltmachzahl von etwa 3 mit Düsendruckverhältnissen zwischen ca. 4 und 40 eingeschaltet ist, gelangt die über den Einlaß 4 einströmende Luft über das dann geöffnete Umschaltorgan 20 in den Turboeinlaßkanal 22, wo sie mittels der Verdichter 18, 24 komprimiert und anschließend in der Brennkammer 26 zusammen mit flüssig gespeichertem Wasserstoff verbrannt wird. Nach der arbeitsleistenden Teilexpansion in der Turbine 28 wird der Heißgasstrom über den ebenfalls geöffneten Ringschieber 30 in den Stauluftkanal 8 eingeleitet und nach Passieren der Einspritzvorrichtungen 12 und der Brennkammer 10, die im Turbobetrieb etwa ab ach 0,9 als Nachbrenner mit Wasserstoff als Treibstoff betrieben wird, in der Düse 16 schaberzeugend entspannt. Zur Absaugung der Rumpfgrenzschicht stromaufwärts des Lufteinlasses 4 im Turbobetrieb ist eine in Fig. 1 in der Schließlage gezeigte Verschlußklappe 34 vorgesehen, an die sich ein zwischen der Flugzeug-Unterseite und dem Hyperschall-Triebwerk zum Triebwerkheck verlaufender Grenzschichtkanal 36 anschließt.In turbo mode, which is switched on from the start up to a switchover number of approximately 3 with nozzle pressure ratios between approximately 4 and 40, the air flowing in via the inlet 4 reaches the turbo inlet duct 22 via the then open switchover element 20, where it is compressed by means of the compressors 18, 24 compressed and then burned in the combustion chamber 26 together with liquid hydrogen. After the partial expansion in the turbine 28, the hot gas flow is introduced into the ram air duct 8 via the ring slide 30, which is also open, and after passing through the injectors 12 and the combustion chamber 10, which is operated in turbo mode from about 0.9 as an afterburner with hydrogen as fuel , relaxed in the nozzle 16 scraping. To extract the fuselage boundary layer upstream of the air inlet 4 in turbo mode, a closure flap 34 is shown in FIG.

Im Staustrahlbetrieb, der von der Umschalt-Machzahl bis zu Hyperschall-Fluggeschwindigkeiten von mehr als 6 Mach mit Düsendruckverhältnissen zwischen etwa 35 und 500 eingeschaltet ist, sind das Umschaltorgan 20 und der Ringschieher 30 ebenso wie die Verschlußklappe 34 geschlossen, so daß das Turbotriebwerk 6 und der Grenzschichtkanal 36 deaktiviert sind, und der gesamte Luftstrom wird einlaufseitig mittels einer verstellbaren Rampenanordnung 38 stoßverdichtet und strömt über den Stauluftkanal 8 in die Brennkammer 10, wo er zusammen mit über die Einspritzvorrichtungen 12 zugeführtem Wasserstoff verbrannt wird und anschließend in dem von der Schubdüse 16 begrenzten Heißgas-Strömungskanal schuberzeugend expandiert.In the ram jet mode, which is switched on from the Mach Mach number to hypersonic flight speeds of more than 6 Mach with nozzle pressure ratios between approximately 35 and 500, the switching member 20 and the ring valve 30 as well as the closure flap 34 are closed, so that the turbo engine 6 and the boundary layer channel 36 is deactivated, and the entire air flow is impact-compressed by means of an adjustable ramp arrangement 38 and flows via the ram air channel 8 into the combustion chamber 10, where it is burned together with hydrogen supplied via the injectors 12 and then expands in the hot gas flow channel delimited by the thrust nozzle 16 to produce thrust.

Die Ausbildung der Schubdüse 16 und die unterschiedlichen Querschnittskonfigurationen des von ihr begrenzten Heißgas-Strömungskanals in den verschiedenen Flugzuständen werden anhand der Fig. 2 bis 4 näher erläutert. Als Hauptbestandteile enthält die Schubdüse 16 eine an die zylindrische Brennkammerwand 40 anschließende, triebwerksfeste, konvergent-divergente, rotationssymmetrische Düsenaussenwand 42, einen pilzförmigen Zentralkörper 44, der auf einem mittleren, triebwerksfesten Träger 46 axial verstellbar angeordnet ist, ein axial verfahrbares, divergentes, ebenfalls rotationssymmetrisches Schubdüsen-Verlängerungsteil 48 und einen die Düsenaußenwand 42 im Bereich des konvergenten Wandabschnittes ringförmig umschliessenden Kaltluft-Verteilerring 50 mit einem an das hintere Ende des Grenzschichtkanals 36 angeschlossenen Luftzulauf 52.The design of the thrust nozzle 16 and the different cross-sectional configurations of the hot gas flow channel delimited by it in the different flight conditions are explained in more detail with reference to FIGS. 2 to 4. The main components of the thrust nozzle 16 include an engine-fixed, convergent-divergent, rotationally symmetrical nozzle outer wall 42 connected to the cylindrical combustion chamber wall 40, a mushroom-shaped central body 44, which is axially adjustable on a central, engine-fixed carrier 46, an axially movable, divergent, also rotationally symmetrical Thrust nozzle extension part 48 and a cold air distributor ring 50, which surrounds the nozzle outer wall 42 in the region of the convergent wall section, with an air inlet 52 connected to the rear end of the boundary layer channel 36.

Im Turbobetrieb ohne Nachverbrennung, d.h. bei Fluggeschwindigkeiten bis zu etwa 0,9 Mach, liegt der Zentralkörper 44 mit seinem größten Querschnitt D zwischen der engsten Stelle und dem Austrittsende 54 der Düsenaußenwand 42 und begrenzt gemeinsam mit dieser einen konvergent-divergenten Heißgas-Strömungskanal 56 mit einem kreisringförmigen Düsenhalsquerschnitt h1, während das Düsenverlängerungsteil 48, das in diesem Flugzustand nicht zur Erweiterung des Heißgas-Strömungskanals 56 benötigt wird, in die eingefahrene Lage zurückgezogen ist und in dieser die äußere Ringwand eines an den Verteilerring 50 anschließenden Luftauslaßkanals 58 bildet, der das divergente Ende der Düsenaußenwand 42 ringförmig umschließt, wobei das Verlängerungsteil 48 in dieser Lage über eine Gleitdichtung 60 mit einem in Richtung des Düsenaustrittsendes offenen Ringspalt 62 des Verteilerrings 50 zusammenwirkt. Auf diese Weise wird die in diesem Zustand einlaufseitig abgetrennte Grenzschichtluft über den zwischen Düsenaußenwand 42 und Verlängerungsteil 48 begrenzten Luftauslaßkanal 58 über den gesamten Außenumfang des Austrittsendes 54 eingeblasen, wodurch nicht nur ein zusätzlicher Schubeffekt erzielt, sondern vor allem auch die Totwasserzone 64, die sich sonst in diesem Flugzustand zwischen dem - in den Figuren gestrichelt dargestellten - Schubstrahl stromabwärts der Düsenaußenwand 42 und der Umgebungsluftströmung (strichpunktiert dargestellt) ausbildet, wirksam abgebaut.In turbo operation without post-combustion, ie at flight speeds of up to about 0.9 Mach, the central body 44 with its largest cross section D lies between the narrowest point and the outlet end 54 of the outer wall 42 of the nozzle and, together with it, delimits a convergent-divergent hot gas flow channel 56 an annular nozzle neck cross section h 1 , while the nozzle extension part 48, which is not required in this flight state for expanding the hot gas flow channel 56, is retracted into the retracted position and in this the outer ring wall of an air outlet channel 58 adjoining the distributor ring 50, which forms the surrounds the divergent end of the nozzle outer wall 42 in an annular manner, the extension part 48 interacting in this position via a sliding seal 60 with an annular gap 62 of the distributor ring 50 which is open in the direction of the nozzle outlet end. In this way, the boundary layer air separated in this state on the inlet side is blown in via the air outlet duct 58 delimited between the outer wall 42 of the nozzle and the extension part 48 over the entire outer circumference of the outlet end 54, which not only achieves an additional thrust effect, but above all also the dead water zone 64, which is otherwise in this flight state between the - shown in dashed lines in the figures - Shear jet downstream of the nozzle outer wall 42 and the ambient air flow (shown in phantom) forms, effectively reduced.

Fig. 3 zeigt die Schubdüse 16 im Turbobetrieb nahe der Unschalt-Machzahl. Der Zentralkörper 44 ist auf dem Träger 46 mit dem größten Zentralkörper-Durchmesser D bis zum Austrittsquerschnitt der festen Düsenaußenwand 42 verstellt, so daß der Düsenhalsquerschnitt auf h2 vergrößert ist. Das Düsenverlängerungsteil 48 befindet sich weiterhin im eingefahrenen Zustand, ist dabei jedoch infolge der Gleitdichtwirkung ohne Dichtverlust gegenüber dem Ringspalt 62 begrenzt axial verstellbar und ermöglicht somit eine Regelung des wirksamen Strömungsquerschnitts des Luftauslaßkanals 58, über den gemäß Fig. 3 weiterhin - nunmehr allerdings über einen wesentlich verkleinerten Strömungsquerschnitt - Grenzschichtluft in die Totwasserzone 64 ausgeblasen wird.3 shows the thrust nozzle 16 in turbo mode close to the Mach Mach number. The central body 44 is adjusted on the carrier 46 with the largest central body diameter D up to the outlet cross section of the fixed nozzle outer wall 42, so that the nozzle neck cross section is increased to h 2 . The nozzle extension part 48 is still in the retracted state, but is limited axially adjustable due to the sliding sealing effect without loss of sealing with respect to the annular gap 62 and thus enables regulation of the effective flow cross section of the air outlet duct 58, via which according to FIG. 3 continues - but now essentially reduced flow cross-section - boundary layer air is blown out into the dead water zone 64.

Fig. 4 zeigt in der oberen Hälfte die Schubdüse 16 kurz nach dem Umschalten auf den Stauluftantrieb. Das Düsenverlängerungsteil 48 ist aus dem mit der Gleitdichtung 60 zusammenwirkenden Einfahrzustand nach hinten in die Ausfahrposition verschoben, in der es formgleich divergent an das Austrittsende 54 der Düsenaußenwand 42 anschließt, wodurch das Erweiterungsverhältnis der Düsenaußenkontur deutlich auf einen der Austrittsfläche A des Verlängerungsteils 48 im Verhältnis zur kleinsten Querschnittsfläche E der Düsenaußenwand 42 entsprechenden Wert erhöht wird. In diesem Flugzustand reicht der Schubstrahl bis zur Umgebungsströmung, so daß die Störwirkung einer Totwasserzone am Austrittsende des Heißgas-Strömungskanals entfällt und der Luftauslaßkanal 58 zur Sekundär-Lufteinblasung nicht mehr benötigt wird. Gleichfalls wird auf die Grenzschichtabtrennung verzichtet, und daher die Verschlußklappe 34 in die in Fig. 1 gezeigte Schließlage verschwenkt. Der Zentralkörper 44 befindet sich in der hinteren Endstellung, in der seine größte Querschnittsfläche D in der Austrittsebene A des divergenten Verlängerungsteils 48 liegt und mithin der Düsenhalsquerschnitt h3 des Heißgas-Strömungskanals 56 seinen Maximalwert erreicht.Fig. 4 shows the thrust nozzle 16 in the upper half shortly after switching to the ram air drive. The nozzle extension part 48 is shifted from the retracted state interacting with the sliding seal 60 to the extended position, in which it adjoins the outlet end 54 of the nozzle outer wall 42 in the same shape, whereby the expansion ratio of the nozzle outer contour clearly to one of the outlet surface A of the extension part 48 in relation to smallest cross-sectional area E of the nozzle outer wall 42 corresponding value is increased. In this flight state, the thrust jet extends to the ambient flow, so that the disturbing effect of a dead water zone at the outlet end of the hot gas flow channel is eliminated and the air outlet channel 58 is no longer required for secondary air injection. Likewise, the boundary layer separation is dispensed with, and therefore the closure flap 34 is pivoted into the closed position shown in FIG. 1. The central body 44 is in the rear end position, in which its largest cross-sectional area D lies in the exit plane A of the divergent extension part 48 and the nozzle neck cross section h 3 of the hot gas flow channel 56 therefore reaches its maximum value.

Bei einer weiteren Erhöhung der Fluggeschwindigkeit bis auf annähernd 7 Mach wird der Zentralkörper 44 unter kontinuierlicher Verringerung des Düsenhalsquerschnitts (und hierzu gegenläufiger Vergrößerung der Heißgas-Austrittsfläche am divergenten Ende des Düsenverlängerungsteils 48) nach vorne verschoben, bis der größte Zentralkörperquerschnitt D in der Ebene E kleinsten Düsenwandaußendurchmessers liegt und mithin der von der Schubdüse begrenzte Heißgas-Strömungskanal 56 seinen minimalen Düsenhalsquerschnitt h4 und das maximale Expansionsverhältnis erreicht (untere Hälfte der Fig. 4; und Fig. 1). Im Hinblick auf eine gesteuerte Schubstrahlexpansion stromabwärts der ausgefahrenen Schubdüsenverlängerung 48 ist eine an diese formgleich anschließende Expansionsrampe 66 zur einseitigen Schubstrahlbegrenzung am Heckende des Flugkörpers 2 angeordnet.With a further increase in the airspeed up to approximately 7 Mach, the central body 44 is moved forward with a continuous reduction in the nozzle neck cross section (and the opposite increase in the hot gas outlet area at the divergent end of the nozzle extension part 48) until the largest central body cross section D in the plane E is smallest Is the outer diameter of the nozzle wall and therefore the hot gas flow channel 56 delimited by the thrust nozzle reaches its minimum nozzle neck cross section h 4 and the maximum expansion ratio (lower half of FIGS. 4; and 1). With regard to a controlled thrust jet expansion downstream of the extended thrust nozzle extension 48, an expansion ramp 66 of the same shape is arranged at the tail end of the missile 2 for one-sided thrust jet limitation.

Claims (10)

  1. An air-breathing jet engine, particularly a hypersonic combination engine, having at least
    - one combustion chamber (10) and a thrust nozzle (16) disposed downstream thereof,
    - a nozzle outer wall (42) fixed to the engine and divergent towards the nozzle outlet end, characterised by
    - on axially displaceable thrust nozzle extension piece (48) of annular and divergent construction, which in its extended position adjoins, with a substantially identical shape, the outlet end (54) of the nozzle outer wall (42) fixed to the engine, and
    - secondary air injection, which can be selectively connected by axially displacing the thrust nozzle extension piece (48), at the outlet end (54) of the nozzle outer wall (42) fixed to the engine, wherein in its retracted state the thrust nozzle extension piece (48) forms an air outlet duct (58), which annularly surrounds the nozzle outer wall (42) fixed to the engine, to the secondary air injection.
  2. A jet engine according to claim 1,
    wherein in its retracted state the thrust nozzle extension piece (48) is restrictedly axially adjustable for varying the effective flow cross-section of the air outlet duct (58).
  3. A jet engine according to claim 1 or 2,
    having a distributor ring (50), which fits round the nozzle outer wall (42) and is fixed to the engine, for the supply of secondary air to the air outlet duct (58), wherein the distributor ring (50) has an annular gap (62) which is open towards the nozzle outlet end and which in the retracted state of the thrust nozzle extension piece (48) is surrounded sealed by the latter.
  4. A jet engine according to claims 2 and 3,
    having a sliding seal (60) between the distributor ring (50) and the thrust nozzle extension piece (48), which sliding seal is effective within the axial range of adjustment in the retracted state of the thrust nozzle extension piece (48).
  5. A jet engine according to claims 3 or 4,
    wherein the distributor ring (50) is disposed in the region of the smallest outside diameter (E) of the nozzle outer wall (42) fixed to the engine.
  6. A jet engine according to any one of the preceding claims,
    having shut-off devices (34) disposed upstream of the air outlet duct (58) for shutting off the secondary air flow in the extended position of the thrust nozzle extension piece (48).
  7. A jet engine according to any one of the preceding claims,
    with boundary layer air separation at the engine inlet for the injection of secondary air in the retracted state of the nozzle extension piece (48).
  8. A jet engine, particularly an air-breathing jet engine according to claim 1, having at least
    - one combustion chamber (10) and a thrust nozzle (16) disposed downstream thereof,
    - a nozzle outer wall (42) fixed to the engine and divergent towards the nozzle outlet end, characterised by
    - an axially displaceable thrust nozzle extension piece (48) of divergent construction, which in its extended position adjoins, with a substantially identical shape, the outlet end (54) of the nozzle outer wall (42) fixed to the engine, and
    - an axially adjustable central body (44) of mushroom-shaped construction, which, in the extended position of the thrust nozzle extension piece (48), can be advanced until its largest central body cross-section (D) is in the region of the thrust nozzle extension piece (48).
  9. A jet engine according to claim 8,
    wherein the largest cross-section (D) of the central body (44) can be adjusted as far as the outlet plane (A) of the extended thrust nozzle extension piece (48).
  10. A jet engine according to any one of the preceding claims,
    having an expansion ramp (66), for delimiting the thrust jet on one side, on the flying object (2) associated with the engine and downstream of the thrust nozzle (16), which expansion ramp adjoins the divergent thrust nozzle extension piece (48) with a shape identical thereto in the extended position of the latter.
EP93108556A 1992-07-11 1993-05-27 Jet engine Expired - Lifetime EP0578951B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE4222947 1992-07-11
DE4222947A DE4222947C2 (en) 1992-07-11 1992-07-11 Jet engine

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EP0578951A1 EP0578951A1 (en) 1994-01-19
EP0578951B1 true EP0578951B1 (en) 1996-07-17

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US5351480A (en) 1994-10-04
JPH06159136A (en) 1994-06-07
DE4222947C2 (en) 1995-02-02
EP0578951A1 (en) 1994-01-19
DE4222947A1 (en) 1994-01-13
JP2767186B2 (en) 1998-06-18

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